The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages that extract energy therefrom. A high pressure turbine first receives the hot combustion gases from the combustor and includes a stator nozzle which directs the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a supporting rotor disk. In a two-stage turbine, a second stage stator nozzle is disposed downstream of the first stage blades followed in turn by a row of second stage rotor blades extending radially outwardly from another supporting disk.
The first and second rotor disks are joined to the compressor by a corresponding rotor shaft for powering the compressor during operation. A multistage low pressure turbine follows the two-stage high pressure turbine and is typically joined by a second rotor shaft to a fan disposed upstream from the compressor in a typical turbofan aircraft engine configuration for powering an aircraft in flight.
As the combustion gases flow downstream through the turbine stages energy is extracted therefrom and the pressure thereof is reduced. A substantial pressure drop occurs across the second stage turbine nozzle, for example, and an interstage seal is typically provided thereat to seal combustor gas leakage around the nozzle.
More specifically, an annular interstage seal ring is mounted axially between the first two rotor disks for rotation therewith during operation, and includes labyrinth seal teeth which extend radially outwardly. A honeycomb stator seal is mounted to the inner end of the second stage nozzle in close proximity to the seal teeth for effecting labyrinth seals therewith for minimizing fluid flow therebetween.
The interstage seal ring includes an annular forward portion which defines a forward cavity on one side of the seal teeth, and an aft portion which defines an aft cavity on the opposite side of the seal teeth.
The nozzle vanes are hollow and provided with a portion of pressurized air from the compressor which is used for cooling the vanes during operation. A portion of the vane air is then channeled radially inwardly through the inner band and discharged through corresponding rows of forward and aft purge holes which supply purge air into the corresponding forward and aft purge cavities on opposite sides of the seal teeth.
In order to enhance cooling of the nozzle vanes themselves, the vanes typically include one or more impingement baffles or inserts therein which have thin sheet metal construction with a multitude of impingement holes therethrough. The surrounding wall of the impingement baffle is spaced closely adjacent to the inner surface of the hollow vanes for discharging corresponding jets of impingement air thereagainst for enhanced cooling thereof. The spent impingement air may then be discharged through various film cooling holes formed through the pressure or suction sides, or both, of the vanes.
The radially outer and inner nozzle bands supporting the vanes provide corresponding boundaries for the combustion gas flow and require correspondingly less cooling thereof. In a typical configuration, each nozzle vane includes an inlet tube or spoolie extending outwardly from the outer band in which cooling air from the compressor is provided. The air travels radially inwardly through the impingement baffle inside each vane, and typically is channeled in part radially through the inner band for providing purge air into the corresponding forward and aft purge cavities.
In one configuration found in commercial service, a transfer tube extends through the inner band for providing pre-impingement air directly into a small cavity created under an individual vane by a sheet metal cover spaced closely adjacent thereto. The sheet metal cover itself is impingement cooled by the air channeled through the transfer tube, with the spent impingement air then being directed into the forward purge cavity, for example.
The interstage honeycomb seal typically includes a sheet metal backing sheet or plate which is suitably fixedly attached to corresponding portions of the inner band, typically without dedicated cooling circuits therefor.
This configuration of the inner band and honeycomb seal attached thereto requires multiple parts which increases the cost and complexity of the configuration. And, this configuration enjoys local cooling capability limited to directly below the individual nozzle vanes.
However, in the development of an improved turbofan gas turbine engine, the combustion gases discharged from the combustor and channeled to the second stage turbine nozzle have a maximum temperature or peak biased closer to the inner band, than the typical center-peaked temperature profile in previous engines. Accordingly, the inner band is subject to a greater heat load during operation and requires a cooling configuration specifically configured therefor for ensuring a suitable useful life of the second stage nozzle during operation.
Accordingly, it is desired to provide an improved turbine nozzle having inner band cooling in a configuration supporting an interstage honeycomb seal.
A turbine nozzle includes hollow nozzle vanes mounted between outer and inner bands. The inner band includes an integral skirt around the perimeter thereof. A backing sheet of a honeycomb seal is mounted to the skirt to define a cavity therein. The inner band includes supply apertures for providing cooling air into the cavity, and the skirt includes forward and aft purge holes for discharging the cooling air therefrom. During operation, the air channeled through the cavity cools the backside of the inner band and is then discharged through the purge holes.